Typically, gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine vane and blade assemblies to these high temperatures. As a result, turbine vanes and blades must be made of materials capable of withstanding such high temperatures. In addition, turbine vanes and blades often contain cooling systems for prolonging the life of the vanes and blades and reducing the likelihood of failure as a result of excessive temperatures.
Typically, turbine vanes are formed from an elongated portion forming a vane having one end configured to be coupled to a vane carrier and an opposite end configured to be movably coupled to an inner endwall. The vane is ordinarily composed of a leading edge, a trailing edge, a suction side, and a pressure side. The inner aspects of most turbine vanes typically contain an intricate maze of cooling circuits forming a cooling system. The cooling circuits in the vanes receive air from the compressor of the turbine engine and pass the air through the ends of the vane adapted to be coupled to the vane carrier. The cooling circuits often include multiple flow paths that are designed to maintain all aspects of the turbine vane at a relatively uniform temperature. At least some of the air passing through these cooling circuits is exhausted through orifices in the leading edge, trailing edge, suction side, and pressure side of the vane.
Many conventional turbine vanes also include film cooling holes in the endwall of the vane. The film cooling holes provide discrete cooling but suffer from numerous drawbacks. For instance, high film cooling effectiveness is difficult to establish and maintain in a highly turbulent environment and large pressure differential region, such as at the intersection between the leading edge and the endwall. In addition, the large pressure gradient that exists at the intersection between the leading edge and the endwall often disrupts the film cooling established by the film cooling holes. Furthermore, the areas between the film cooling orifices and areas immediately downstream from the film cooling orifices are typically not in contact with the cooling fluids and therefore are not cooled by the cooling fluids. Consequently, these areas are more susceptible to thermal degradation and over temperatures.
As shown in FIG. 1, turbine vanes often experience horseshoe vortex flow phenomenon created by the combination of hot gas radial velocity and static pressure gradient forces at the intersection of the airfoil leading edge and the endwall. As the hot gas flow encounters the airfoil and collides with the leading edge, the horseshoe vortex separates into pressure side and suction side downward vortices. Initially, the pressure vortex sweeps downward and flows along the airfoil pressure side. However, the pressure side vortex shifts to the suction side of an adjacent airfoil due to the pressure differential between the pressure side and suction side of the adjacent airfoil. As the vortex flows to the suction side, the vortex grows in size and strength and becomes much larger than the vortex located at the suction side and creates a high heat transfer and high gas temperature region at the suction side.
Conventional backside impingement has not been successful in cooling this region. In addition, traditional film cooling has likewise been unsuccessful because effective cooling may only be partially achieved when the impingement orifices are tightly packed together. However, such formation of closely packed film cooling orifices is difficult to manufacture. Conversely, spacing the film cooling orifices further apart creates regions that do not receive film cooling air and are more susceptible to thermal degradation. Thus, such configuration is not an acceptable alternative. Thus, a need exists for a turbine vane having increased cooling efficiency for dissipating heat at the intersection of the turbine blade and the endwall.